It is well known that nickel based superalloys are extensively employed in high performance environments. Such alloys have been used extensively in jet engines and in gas turbines where they must retain high strength and other desirable physical properties at elevated temperatures of a 1000F. or more.
The strength of these alloys is related to the presence of a strengthening precipitate, which in many cases is a .gamma.' precipitate or .gamma." precipitate. More detailed characteristics of the phase chemistry of precipitates are given in "Phase Chemistries in Precipitation-Strengthening Superalloy" by E. L. Hall, Y. M. Kouh, and K. M. Chang [Proceedings of 41st. Annual Meeting of Electron Microscopy Society of America, August 1983 (p. 248)].
The following U.S. patents disclose various nickel-base alloy compositions, some of which contain such precipitates: U.S. Pat. Nos. 2,570,193; 2,621,122; 3,046,108; 3,061,426; 3,151,881; 3,166,412; 3,322,534; 3,343,950; 3,575,734; 3,576,681; 4,207,098 and 4,336,312. The aforementioned patents are representative of the many alloying situations reported to date in which many of the same elements are combined to achieve distinctly different functional relationships between the elements such that phases from providing the alloy system with different physical and mechanical characteristics are formed. Nevertheless, despite the large amount of data available concerning the nickel-base alloys, it is still not possible for workers in the art to predict with any degree of accuracy the physical and mechanical properties that will be displayed by certain concentrations of known elements used in combination to form such alloys even though such combination may fall within broad, generalized teachings in the art, particularly when the alloys are processed using heat treatments different from those previously employed.
It is known that some of the most demanding sets of properties for superalloys are those which are needed in connection with jet engine construction. Of the sets of properties which are needed those which are needed for the moving parts of the engine are usually greater than those needed for static parts although the sets of needed properties are different for the different components of an engine.
Because some sets of properties have not been attainable in cast alloy materials, resort is sometimes had to the preparation of parts by powder metallurgy techniques. However, one of the limitations which attends the use of powder metallurgy techniques in preparing moving parts for jet engines is that of the purity of the powder. If the powder contains impurities such as a speck of ceramic or oxide the place where that speck occurs in the moving part becomes a latent weak spot where a crack may initiate or a latent crack.
To avoid problems with impure powder and similar problems it is sometimes preferred to form moving parts of jet engines such as disks with alloys which can be cast and wrought.
A problem which has been recognized to a greater and greater degree with many such nickel based superalloys is that they are subject to formation of cracks or incipient cracks, either in fabrication or in use, and that the cracks can actually initiate or propagate or grow while under stress as during use of the alloys in such structures as gas turbines and jet engines. The propagation or enlargement of cracks can lead to part fracture or other failure. The consequence of the failure of the moving mechanical part due to crack formation and propagation is well understood. In jet engines it can be particularly hazardous.
However, what has been poorly understood until recent studies were conducted was that the formation and the propagation of cracks in structures formed of superalloys is not a monolithic phenomena in which all cracks are formed and propagated by the same mechanism and at the same rate and according to the same parameters and criteria. By contrast the complexity of the crack generation and propagation and of the crack phenomena generally, and the interdependence of such propagation with the manner in which stress is applied, is a subject on which important new information has been gathered in recent years. The period during which stress is applied to a member to develop or propagate a crack, the intensity of the stress applied, the rate of application and of removal of stress to and from the member and the schedule of the application was not well understood in the industry until a study was conducted under contract to the National Aeronautics and Space Administration. This study is reported to a technical report identified as NASA CR-165123 issued from the National Aeronautics and Space Administration in August 1980, identified as "Evaluation of the Cyclic Behavior of Aircraft Turbine Disk Alloys", Part II, Final Report, by B. A. Cowles, J. R. Warren and F. K. Hauke, and prepared for the National Aeronautics and Space Administration, NASA Lewis Research Center, Contract NAS3-21379.
A principal unique finding of the NASA sponsored study was that the rate of propagation based on fatigue phenomena or in other words the rate of fatigue crack propagation (FCP) was not uniform for all stresses applied nor to all manners of applications of stress. More importantly, the finding was that fatigue crack propagation actually varied with the frequency of the application of stress to the member where the stress was applied in a manner to enlarge the crack. More surprising still, was the finding from the NASA sponsored study that the application of stress of lower frequencies rather than at the higher frequencies previously employed in studies actually increased the rate of crack propagation. In other words the NASA study revealed that there was a time dependence in fatigue crack propagation. Further the time dependence of fatigue crack propagation was found to depend not on frequency alone but on the time during which the member was held under stress or a so-called hold-time.
Following the discovery of this unusual and unexpected phenomena of increased fatigue crack propagation at lower stress frequencies there was some belief in the industry that this newly discovered phenomena represented an ultimate limitation on the ability of the nickel based superalloys to be employed in the stress bearing parts of the turbines and aircraft engines and that all design effort had to be made to design around this problem.
However, it has been discovered that it is feasible to construct parts of nickel based superalloys for use at high stress in turbines and aircraft engines with greatly reduced crack propagation rates.
The development of the superalloy compositions and methods of their processing of this invention focuses on the fatigue property and addresses in particular the time dependence of crack growth.
Crack growth, i.e., the crack propagation rate, in high-strength alloy bodies is known to depend upon the applied stress (.sigma.) as well as the crack length (a). These two factors are combined by fracture mechanics to form one single crack growth driving force; namely, stress intensity K, which is proportional to .sigma..sqroot.a. Under the fatigue condition, the stress intensity in a fatigue cycle represents the maximum variation of cyclic stress intensity (.DELTA.K), i.e., the difference between K.sub.max and K.sub.min. At moderate temperatures, crack growth is determined primarily by the cyclic stress intensity (.DELTA.K) until the static fracture toughness K.sub.IC is reached. Crack growth rate is expressed mathematically as da/dN.varies.(.DELTA.K).sup.n. N represents the number of cycles and n is a constant which is between 2 and 4. The cyclic frequency and the shape of the waveform are the important parameters determining the crack growth rate. For a given cyclic stress intensity, a slower cyclic frequency can result in a faster crack growth rate. This undesirable time-dependent behavior of fatigue crack propagation can occur in most existing high strength superalloys. According to this hold time pattern, the stress is held for a designated hold time each time the stress reaches a maximum in following the normal sine curve. This hold time pattern of application of stress is a separate criteria for studying crack growth. This type of hold time pattern was used in the NASA study referred to above.
The design objective is to make the value of da/dN as small and as free of time-dependency as possible.
It is pointed out in copending application Ser. No. 907,550, filed Sept. 15, 1986 that time dependent fatigue crack propagation can be reduced significantly by a thermal treatment of .gamma.' strengthened nickel base superalloys which have more than 35 volume percent of strengthening precipitate. As is pointed out in this copending application, the method involves a high temperature solutioning (supersolvus) solutioning of the .gamma.' precipitate followed by a controlled cobling at less than 250.degree. F. per minute.
However, it has been found that the method of copending application Ser. No. 907,550 does not yield the beneficial results taught in that application when the method is applied to alloys with low precipitate content. For example, the method does not produce the fatigue crack propagation reduction when applied to Waspalloy or to IN 718 alloy. Waspalloy is .gamma.' hardened and has less than 35 volume percent and preferably about 30 volume percent .gamma.' precipitate. IN 718 is mainly .gamma." hardened and has less than 35 volume percent and preferably about 20 percent by volume of .gamma.' precipitate.
I have done extensive studies on alloys of such lower .GAMMA.' or .gamma." precipitate content and have heat treated these alloys according to a variety of schedules which restrict fatigue crack propagation properties of alloys having higher precipitate content but without significant beneficial effect. I have found that none of these heat treatments develop different or advantageous microstructures or result in any significant reduction in fatigue crack propagation.
Pursuant to the present invention a method for processing a superalloy containing a lower concentration of strengthening precipitate is provided to produce materials with a superior set or combination of properties for use in advanced engine disk applications. The properties which are conventionally needed for materials used in disk applications include high tensile strength and high stress rupture strength. In addition the alloy prepared by the methods of the subject invention exhibits a desirable property of resisting crack growth propagation. Such ability to resist crack growth is essential for the component low cycle fatigue life or LCF.
In addition to this superior set of properties as outlined above, the alloy processed by the method of the present invention displays good forgeability and such forgeability permits greater flexibility in the use of various manufacturing processes needed in formation of parts such as disks for jet engines.
Superalloys with lower ranges of precipitate content generally have good forgeability and can be subjected to thermomechanical processing. The difference of certain thermomechanical processings on mechanical properties, like strength and rupture life, are known to a degree. However, nothing was known heretofore of the influence, if any, of thermomechanical processings on time-dependent fatigue crack propagation or the rates of such propagation.
As alloy products for use in turbines and jet engines have developed it has become apparent that different sets of properties are needed for parts which are employed in different parts of the engine or turbine. For jet engines the material requirements of more advanced aircraft engines continue to become more strict as the performance requirements of the aircraft engines are increased. The different requirements are evidenced, for example, by the fact that many blade alloys display very good high temperature properties in the cast form. However, the direct conversion of cast blade alloys into disk alloys is very unlikely because blade alloys display inadequate strength at intermediate temperatures of about 700.degree. C. Further, the blade alloys have been found very difficult to forge and forging has been found desirable in the fabrication of blades from disk alloys. Moreover, the crack growth resistance of disk alloys has not been evaluated.
Accordingly, to achieve increased engine efficiency and greater performance, constant demands are made for improvements in the strength and temperature capability of disk alloys as a special group of alloys for use in aircraft engines. Now these capabilities must be coupled with low fatigue crack propagation rates and a low order of time-dependency of such rates.
What was sought in undertaking the work which lead to the present invention was the development of a processing for a disk alloy which resulted in a low or minimum time dependence of fatigue crack propagation and moreover a high resistance to fatigue cracking.